The present invention relates to the general field of aviation turbine engines, and more precisely to the field of fan blade platforms attached to an aviation turbine engine.
In a turbine engine, attached fan blade platforms need to perform several functions. From an aerodynamic point of view, these platforms serve firstly to define the air flow passage. In addition, they must also be capable of withstanding large forces without deforming and of remaining secured to the disk that carries them.
In order to satisfy these various requirements, certain configurations have been proposed in which platforms possess a first portion serving to define the air flow passage and to retain the platform while the engine is rotating, and a second portion serving to limit deformation to the first portion under the effect of centrifugal forces and to hold the platform in position when the engine is stopped.
In existing solutions, the platform may be in the form of a box section beam with a two-dimensional passage wall that is held downstream by a drum and upstream by a shroud, with upstream retention by the shroud taking place above the tooth of the fan disk (a flange of the shroud serves to block the platform axially and radially at its upstream end).
Such retention taking place above the tooth of the disk by using a shroud presents the drawback of imposing a large hub ratio, where the hub ratio is the ratio of the radius measured between the axis of rotation and the point of the leading edge of the blade flush with the surface of the platform divided by the radius measured between the axis of rotation and the outermost point of the leading edge.
In order to optimize the performance of the fan, and more generally of the engine, it is desirable to have available a rotary assembly for an aviation turbine engine that includes an attached fan blade platform that is mounted on a fan disk and that presents a hub ratio that is as small as possible.